In conventional gas turbine engine rotor assemblies, a plurality of aerodynamically shaped blade airfoils are disposed in the flowpath to react with the working fluid of the engine. For example, in a turbofan engine, air passing through the inlet cowling is compressed initially in a multistage fan or low pressure compressor (LPC). A large portion of the air is channeled aft through a duct, bypassing the core engine flowpath. The remainder of the compressed air passes through a multistage high pressure compressor (HPC) where it is further compressed before being mixed with fuel and ignited in the combustion section of the engine. The hot gases subsequently pass through a high pressure turbine (HPT), which is operably connected to the HPC by a shaft, where energy is extracted to drive the HPC. The flow is then directed through a low pressure turbine (LPT), which is operably connected to the LPC by a second shaft, where additional energy is extracted to drive the LPC. The flow is then combined with the bypass flow and exhausted through a nozzle, to provide propulsive thrust to an airframe.
In one conventional style of axial fan and compressor rotor, each stage is comprised of a plurality of removable blades, typically retained in the rim of a rotor disk by means of convoluted blade dovetails and complementary shaped axial slots formed between adjacent disk posts. The interlocking dovetail and slot contours are precisely configured and machined to ensure proper fit and retention of the blades under very high disk rotational speeds and induced centrifugal and aerodynamic loads. Further, the load path through the blade dovetail into the disk must be carefully controlled so as to prevent detrimental vibratory modes as well as avoid excessive component stresses. Failures which occur at the dovetail/disk interface can result in release of the blades from the disk at high rotational speeds resulting in significant secondary damage to the engine.
A fundamental problem associated with gas turbine rotors of this design is the detrimental effect of high pressure cyclic loading through the dovetail and disk post interfaces, commonly referred to as pressure faces. During periods of change in rotor speed such as runup and coastdown, changes in disk rim elastic strain result in relative movement and slippage between the dovetail and disk posts along the pressure faces. This sliding action damages the surfaces by introducing microcracks, prime initiation sites for fatigue cracks, which propagate under the combined effect of high cycle vibration and recurring stress cycling. This condition, commonly referred to as fretting, results in a significant reduction in component fatigue life. Titanium alloys, which are used extensively in modern gas turbine fan and compressor rotor stages due in part to exceptional strength to weight ratios, have been shown to be particularly susceptible to fretting damage, visually apparent as localized zones of surface discoloration. Fretting is most apparent when both the blades and disks are comprised of titanium alloy. In practice, reduction in component fatigue strength of up to 75% and related shortening of component life is common. Fretting tends to be more pronounced in front end blading of fan and compressor rotors, where relative slippage and blade loads are greatest. Integrally bladed disks, commonly referred to as blisks, are sometimes employed in these locations; however, since front end stages are most susceptible to ingested foreign object damage, replaceable blading is desirable for economic considerations. While titanium exhibits particular susceptibility to fretting damage, other conventional steel and nickel based alloys employed in rotors can exhibit similar distress.
Attempts have been made, both in the design and manufacture of rotor disks and blades by those skilled in the art, to either delay the onset of fretting damage or minimize its effect. For example, pressure face contact angle is predetermined to control the amount of relative slippage and the magnitude and direction of the transmitted load. Also, during manufacture, after being ground to precise contour and dimensions, blade dovetails may be shotpeened. While providing compressive surface stresses which make the dovetail pressure face surface less susceptible to microcracking, shotpeening increases average surface roughness to the range of 32 microinches rms or greater, thereby increasing resultant sliding friction. Similarly, disk dovetail slots may be shotpeened after being broached, although this procedure is more difficult and more variable due to limited accessibility. To provide additional margin against the initiation of fretting damage, protective alloy wear coatings such as thin sacrificial layers of copper, nickel and indium, in the range of several mils of thickness, are routinely applied to blade dovetails. Further, lubricants such as molybdenum disulfide are applied to the pressure faces to reduce friction; however, their effectiveness is short lived and periodic reapplication is necessary, costly and inconvenient. These coatings and lubricants can also be employed in cooperation with shim elements located between the dovetails and the disk posts. Also, dovetails slot contours can be modified by undercutting the disk posts to remove material subject to peak surface stresses. While actions such as these are helpful in delaying the onset of fretting, they fail to address the inherent problem of high cycle, high pressure induced sliding damage. With time, relative movement degrades the treated contact surfaces until the dovetail and/or the disk post suffer irreparable fretting damage and fatigue cracking at which point they must be removed from service. Due to the significant secondary damage caused by rotor component failures, the risk associated with operating potentially damaged components is typically unwarranted; therefore, frequent stripping, inspection, and recoating of blade dovetails and ultimately removal and replacement of nondiscrepant components at considerable inconvenience and cost is often required.